1. Field of the Invention
The present invention relates to rocket and ramjet fuel compositions, and more particularly to fuel compositions suitable for use in so-called "hybrid" propulsion rocket engines.
2. Description of the Related Art
As a result of intensive research efforts, rocket engines have undergone significant evolution since the first aerospace vehicles were launched several decades ago. Practitioners have progressively improved engine designs to achieve better performance and safety characteristics, and have also succeeded in simplifying engine constructions for ease of production. Traditional solid-propellant engines, for example, pose safety hazards both in manufacture and operation, owing to the intimate bulk combination of the fuel composition and an oxidant to promote its combustion. Such engines also suffer from performance limitations, and their propellant systems often contain ingredients that contribute to atmospheric pollution.
Liquid-propellant devices, in which fuel and oxidant are separately stored as liquids and sprayed simultaneously into a combustion chamber, can deliver improved performance over solid-propellant engines. However, due in part to the need for precise control over the rate at which fuel and oxidant are introduced, liquid-propellant rocket engines are highly complex to produce and also to maintain. Their intricacy increases the potential for malfunction and safety hazard.
A recent improvement over these traditional solid- and liquid-propellant approaches is the "hybrid" engine, which utilizes both solid and liquid (or gaseous) components. Typically the solid component is a polymeric fuel, and is used in conjunction with a liquid or gaseous oxidizing component (most frequently a liquid that may be introduced as an easily vaporized spray); this model will be assumed for purposes of discussion. Hybrid designs offer numerous advantages. Properly engineered hybrid rocket engines, though simple in construction, are capable of delivering high thrust levels. Separation of fuel and oxidant components promotes safety, and the need to convey only one component into the combustion chamber reduces the regulatory and conduction hardware necessary for operation.
FIG. 1 schematically illustrates the hybrid concept, which includes a pressure casing or shell 10 (fabricated from, for example, a graphite/epoxy composite) that terminates in a nozzle 12, through which exhaust gases are ejected to provide thrust. Within casing 10 and generally conforming to its interior dimension is a continuous cake of solid fuel 14, which is hollowed out to define a combustion chamber or conduit 16 where burning takes place.
A source 20 of oxidant, which may be, for example, liquid oxygen, is introduced into combustion chamber 16 by means of an injector (not shown). The flow of oxidant is controlled by a valve 22. Combustion proceeds by vaporization of the solid fuel and, if a liquid oxidant is used, vaporization of that component as well. A mixture of the vapor-phase fuel and oxidant combusts near the surface 26 of solid fuel cake 14, and the gas flow over solid surface 26 develops a boundary layer with a velocity and temperature distribution that combine to transfer heat to the solid fuel, promoting further vaporization to continue combustion. The overall rate of the combustion process is limited by the rate at which heat is transferred to the solid fuel, since it is heat transfer that determines the vaporization rate.
These phenomena are illustrated diagrammatically in FIG. 2. The rate ds/dt at which solid fuel is consumed (and which, assuming an annular fuel cross-section, is proportional to the rate at which the diameter of combustion chamber 16 linearly increases) depends on the heat-transfer rate dQ/dt into solid fuel cake 14. The heat-transfer rate, in turn, is determined by the velocity and temperature profiles of the boundary layer along surface 26. As shown in the figure, the velocity profile v (where the ordinate represents increasing distance D from solid fuel cake 14 and rightward movement along the abscissa corresponds to increasing velocity) follows a standard turbulent flow pattern, with gases nearly stagnant close to surface 26. The temperature profile T (where rightward movement along the abscissa corresponds to increasing temperature) reaches a maximum at a characteristic value of D. The flow of fuel vapor from the solid surface serves to increase this distance and to increase the thickness of the velocity boundary layer, with the effect of constraining the rate of heat transfer dQ/dt into the solid fuel cake 14.
One clear disadvantage of the hybrid concept results from the low linear burning rate ds/dt inherent to vaporization of a simple polymeric fuel in a combusting boundary-layer flow, which constrains the maximum achievable thrust. Rocket engine thrust is proportional to the product of the exhaust mass flow rate and the exhaust velocity. The latter quantity is largely defined by the chemical compositions of the fuel and the oxidant, and reaches a maximum value at a characteristic ratio of fuel to oxidant. The mass flow rate is the sum of oxidant and fuel flows and, for a fixed ratio of fuel to oxidant, is determined by the mass burning rate of the fuel. In a hybrid rocket this mass burning rate is proportional to the product of fuel density, the exposed surface area 26, and the linear burning rate ds/dt.
The fuel density is fixed by its chemical composition. Therefore, improving the mass flow rate to increase thrust requires expansion of the exposed fuel surface area and/or the linear burning rate. Because conventional polymeric hybrid fuels burn at relatively low linear rates, engineering efforts have focused on ways of increasing the surface area. For example, one current high-thrust hybrid rocket development effort utilizes a "wagon-wheel" fuel grain design with more than 12 axial ports to obtain adequate burning surface area. Kniffen, R. J., "Hybrid Rocket Development at the American Rocket Company," 26th Joint Propulsion Conference, AIAA 90-2762, July 1990. Such complex solid fuel geometries impose high fabrication costs, large inert rocket-engine mass, and overall propulsion-system performance levels that fall well below theoretical limits.